Combustor burner arrangement

ABSTRACT

A burner for a combustor having a burner axis, a fuel lance, an igniter and a main air flow passage(s). The passage is angled relative to the burner axis and creates a main vortex thereabout in a first rotational direction. The main vortex travels in a direction along the burner axis, away from the surface. The fuel lance is located downstream of the igniter with respect to the first rotational direction of the main vortex so that a part of a main air flow washes over the fuel lance and over the igniter. The fuel lance has a fuel lance axis, a liquid fuel tip having a fuel outlet and an array of air assist passages having outlets arranged thereabout. The air assist passages are angled relative to create an air assist vortex about the fuel lance axis in the same rotational direction to the first rotational direction.

FIELD OF INVENTION

The present invention relates to combustion equipment of a gas turbineengine and in particular a burner arrangement of the combustionequipment.

BACKGROUND OF INVENTION

Gas turbines including dry low emission combustor systems can havedifficulty lighting and performing over a full load range when usingliquid fuels. Often this can be because of fuel placement and subsequentatomization of the fuel in mixing air flows particularly at low loadsdemanded from the engine. Ideally, the fuel droplets need to be verysmall and injected into an appropriate part of the airflow entering thecombustor's pre-chamber via an annular array of main air flow swirlersin the vicinity of a burner arrangement to burn in the correct flamelocation. Also the fuel droplets should not contact any wall surface butat the same time the fuel droplets need to come close enough to theigniter so that the igniter can ignite the vaporised fuel on start up.If the fuel droplets contact a surface this can lead to carbon depositsbuilding up or lacquers forming and which can alter airflowcharacteristics or even block air and/or fuel supply holes.

The liquid pilot injection lance can have additional air assistance toaid atomisation of the liquid fuel over a range of fuel flows. This airassistance can be a supplied via a number of air outlets completelysurrounding a fuel orifice or filmer. These air assist outlets areangled to create a pilot fuel/air vortex that rotates in an oppositedirection to the direction of rotation of a main fuel/air vortex. Thisliquid pilot injection lance is in a region prone to liquid fuel contactand as a result tends to incur carbon deposits. The residence time ofthe liquid fuel droplets is a crucial parameter for the fuel droplets tostay close to the surface of the wall of the burner. The longer the timethe liquid fuel droplets reside close to the surface of the wall, thericher the fuel/air mixture and hence there is an increase in carbondeposition, nitrous-oxide (NOx) emissions and higher local heating ofthe surface near the pilot burner which in turn increases thermalgradients that lead to cracking in the surface.

These carbon deposits block the air assistance holes and subsequentlyprevent successful atomisation of the fuel. Poor atomisation of thepilot fuel also causes problems with ignition of the fuel at start-up.This is a common fault with gas turbine fuel injection systems andcarbon build up is a common problem. Consequently, liquid pilotinjection lances are regularly replaced and are considered a consumablepart. This is undesirable because such replacement is expensive, causesthe gas turbine to be off-line halting supply of electricity or powerfor example, and can be unscheduled.

SUMMARY OF INVENTION

One objective of the present invention is to prevent carbon depositsforming on components. Another objective is to prevent carbon depositsforming on a fuel lance of a combustor. Another object is to improve thereliability of igniting the fuel in a combustor. Another objective is toimprove the entrainment of fuel droplets in an air flow. Anotherobjective is to improve the atomisation of liquid fuel in a combustor.Another objective is to prevent liquid fuel contacting a surface withinthe combustor. Another objective is to reduce or prevent scheduled orunscheduled shut down of the engine for maintenance attributed toreplacing or cleaning combustor components subject to carbon depositsand particularly the liquid fuel lance. Yet another objective is toreduce the residence time of the liquid fuel droplets close to theburner surface. A further objective is to reduce high local heating ofthe burner surface. Still further, an objective is to reduce emissionsof the combustor.

These advantages and objectives are realised by the provision of aburner for a combustor of a gas turbine combustor, the burner comprisesa body having a surface and an burner axis, a fuel lance, an igniter anda main air flow passage or passages, the main air flow passage orpassages is angled relative to the burner axis and creates a main vortexabout the burner axis in a first rotational direction, the main vortextravels in a direction along the burner axis and away from the surface,the fuel lance is located downstream of the igniter with respect to thefirst rotational direction of the main vortex so that a part of a mainair flow washes over the fuel lance and then over the igniter, the fuellance comprises a fuel lance axis, a liquid fuel tip having a fueloutlet and an array of air assist passages having outlets arranged aboutthe fuel outlet, wherein the air assist passages are angled relative tothe fuel lance axis to create an air assist vortex about the fuel lanceaxis in the same rotational direction to the first rotational direction.

The main air flow passage or passages may be tangentially angledrelative to the burner axis.

The air assist passages may be radially angled relative to the fuellance axis.

The air assist passages may be radially angled between and including 15°and 60° relative to the fuel lance axis.

The air assist passages may be radially angled approximately 30°relative to a tangent of the fuel lance axis.

The air assist passages may have a tangential angle between +/−45°relative to a tangent of the fuel lance axis.

The air assist passages may have a tangential angle approximately 0°relative to a tangent of the fuel lance axis.

The fuel lance and the igniter may be located at the same radialdistance from the burner axis.

The fuel lance and the igniter may be located at different radialdistances from the burner axis and preferably the igniter is radiallyinwardly of the fuel lance.

The fuel outlet of the fuel lance may be located at or near to thesurface.

The igniter may be at least partly housed within the body and has an endface, the end face is located at or near to the surface.

The burner may comprise an annular array of swirl vanes arranged aboutthe burner axis and which form the main air flow passages.

The main air flow passages may be angled in an anti-clockwise directionand the air assist passages are angled in an anti-clockwise directionrelative to a view on to the surface.

The main air flow passages may be angled in a clockwise direction andthe air assist passages are angled in a clockwise direction relative toa view on to the surface.

In one example, the fuel outlet is a fuel prefilmer and which may bedivergent towards its end and can create a cone of fuel. In anotherexample, the fuel outlet is an orifice and which can create a spray offuel. In yet another example, the fuel outlet is a number of orificesand each orifice can create a spray of fuel.

BRIEF DESCRIPTION OF THE DRAWINGS

Further features, properties and advantages of the present inventionwill become clear from the following description of embodiments inconjunction with the accompanying drawings in which;

FIG. 1 shows part of a turbine engine in a sectional view in which thepresent invention is incorporated,

FIG. 2 shows a perspective schematic view of a section of a combustorunit of turbine engine and in detail a burner arrangement including apilot burner surrounded by a main burner, the pilot burner having aliquid fuel lance and an igniter and is in accordance with presentinvention,

FIG. 3 shows a schematic perspective and cut-away view of part of thepilot burner and in detail the liquid fuel lance in accordance withpresent invention,

FIG. 4 is a view along a combustor axis and onto the surface of theburner as shown in FIG. 2 and where the pilot burner is generallysurrounded by the main burner having an annular array of swirler vanes,the pilot burner having a liquid fuel lance in accordance with presentinvention,

FIG. 5 and FIG. 6 show sectional views of the main air flow along pathsA-A and B-B respectively as shown in FIG. 4 and illustrates respectivedistributions of fuel droplets issuing from the liquid fuel lance,

FIG. 7 is a view on a tip of an embodiment of the present liquid fuellance and generally along its axis showing an array of outlets arrangedaround a fuel outlet; the array of outlets directs a pilot air flow toimpinge on, shearing and atomizing a liquid fuel film,

FIG. 8A is an isometric view of the liquid fuel lance schematicallyshowing the relative radial angle μ of one of the air passages; theother air passages have been omitted for clarity,

FIG. 8B is a view on the exposed surface of the tip 72 of the liquidfuel lance as seen in FIG. 3,

FIG. 9 is a view on the surface of the burner and along the burner'scentral axis and indicates the orientation of the liquid fuel lancerelative to the main air flow from the main burner and relative to theburner's central axis and in accordance with the present invention,

DETAILED DESCRIPTION OF INVENTION

FIG. 1 shows an example of a gas turbine engine 10 in a sectional viewand generally arranged about a longitudinal axis 20. The gas turbineengine 10 comprises, in flow series, an inlet 12, a compressor section14, a combustor section 16 and a turbine section 18 which are generallyarranged in flow series and generally in the direction of thelongitudinal or rotational axis 20. The gas turbine engine 10 furthercomprises a shaft 22 which is rotatable about the rotational axis 20 andwhich extends longitudinally through the gas turbine engine 10. Theshaft 22 drivingly connects the turbine section 18 to the compressorsection 12. The combustor section 16 comprises an annular array ofcombustor units 16 only one of which is shown.

In operation of the gas turbine engine 10, air 24, which is taken inthrough the air inlet 12 is compressed by the compressor section 14 anddelivered to the combustion section or unit 16. The combustor unit 16comprises a burner plenum 26, a pre-chamber 29, a combustion chamber 28defined by a double walled can 27 and at least one burner 30 fixed toeach combustion chamber 28. The pre-chamber 29, the combustion chamber28 and the burner 30 are located inside the burner plenum 26. Thecompressed air 31 passing through the compressor 12 enters a diffuser 32and is discharged from the diffuser 32 into the burner plenum 26 fromwhere a portion of the air enters the burner 30 and is mixed with agaseous and/or liquid fuel. The air/fuel mixture is then burned and theresulting combustion gas 34 or working gas from the combustion chamberis channeled via a transition duct 35 to the turbine section 18.

The turbine section 18 comprises a number of blade carrying rotor discs36 attached to the shaft 22. In the present example, two discs 36 eachcarry an annular array of turbine blades 38. However, the number ofblade carrying rotor discs could be different, i.e. only one disc ormore than two rotor discs. In addition, guiding vanes 40, which arefixed to a stator 42 of the gas turbine engine 10, are disposed betweenthe turbine blades 38. Between the exit of the combustion chamber 28 andthe leading turbine blades 38 inlet guiding vanes 44 are provided.

The combustion gas 34 from the combustion chamber 28 enters the turbinesection 18 and drives the turbine blades 38 which in turn rotates theshaft 22 to drive the compressor section 12. The guiding vanes 40, 44serve to optimise the angle of the combustion or working gas on to theturbine blades 38. The compressor section 12 comprises an axial seriesof guide vane stages 46 and rotor blade stages 48.

The terms upstream and downstream refer to the flow direction of theairflow and/or working gas flow through the engine unless otherwisestated. The terms forward and rearward refer to the general flow of gasthrough the engine. The terms axial, radial and circumferential are madewith reference to the rotational axis 20 of the engine unless otherwisestated.

FIG. 2 is a perspective view of a part of the combustor 16 showing theburner 30, the pre-chamber 29 and part of the combustion chamber 28. Thecombustion chamber 28 is formed with a tubular-like shape by the doublewalled can 27 (shown in FIG. 1) having and extending along a combustoraxis 50. The combustor 16 extends along the combustor axial 50 andcomprises the pre-chamber 29 and the main combustion chamber 28, whereinthe latter extends in a circumferential direction 61 around thecombustor axis 50 and generally downstream, with respect to the gas flowdirection, of the pre-chamber volume 29.

The burner 30 comprises a pilot burner 52 and a main burner 54. Thepilot burner 52 comprises a burner body 53, a liquid fuel lance 56 andan igniter 58. The main burner 54 comprises a swirler arrangement 55having an annular array of swirler vanes 60 defining passages 62therebetween. The annular array of swirler vanes 60 are arrangedgenerally about a burner axis 50, which in this example is coincidentwith the combustor axis 50, and in conventional manner. The swirlerarrangement 55 includes main fuel injection ports which are not shown,but are well known in the art. The main burner 54 defines part of thepre-chamber 29. The pilot burner 52 is located in an aperture 57 andgenerally radially inwardly, with respect to the burner/combustor axis50, of the main burner 54. The pilot burner 52 has a surface 64 thatdefines part of an end wall of the pre-chamber 29. The end wall isfurther defined by the main burner 54.

The liquid fuel lance 56 is at least partly housed in a first hole 66defined in the burner body 53 of the pilot burner 52. A pilot air flowpassage 69 is formed between the liquid fuel lance 56 and the walls ofthe first hole 66. The liquid fuel lance 56 comprises an elongate fuellance body 86 and a liquid fuel tip 72. The elongate fuel lance body 86is generally cylindrical and defines a fuel flow passage 70. The liquidfuel tip 72 is mounted at one end of the elongate fuel lance body 86 andis located near to or at the surface 64. The liquid fuel lance 56 willbe described in more detail with reference to FIG. 3. The igniter 58 ishoused in a second passage 74 defined in the burner body 53 of the pilotburner 52. The end of the igniter 58 is located near to or at thesurface 64. The igniter 58 is a well known device in the art and thatrequires no further description. In other combustors 16 it is possiblethat more than one liquid fuel lance and/or more than one igniter may beprovided.

During operation of the gas turbine engine and more particularly atengine start-up, a starter-motor cranks the engine such that thecompressor 14 and turbine 16 are rotated along with the shaft 22. Thecompressor 14 produces a flow of compressed air 34 which is delivered toone or more of the combustor units 16. A first or major portion of thecompressed air 34 is a main air flow 34A which is forced through thepassages 62 of the swirler arrangement 55 where the swirler vanes 60impart a swirl to the compressed air 34 as shown by the arrows. A secondor minor portion of the compressed air 31 is a pilot air flow 34B whichis forced through the pilot air flow passage 69. The pilot air flow 34Bcan also be referred to as an air assistance flow. Liquid fuel 76 isforced through the fuel flow passage 70 and is mixed with the pilot airflow 34B and the main air flow 34A in order to atomise the liquid fuel.Atomisation of the liquid fuel into very small droplets increasessurface area to enhance subsequent vaporisation.

The main air flow 34A generally swirls around the combustor axis 50. Theswirler vanes 60 impart a tangential direction component to the main airflow 34A to cause the bulk main air flow 34 to have a circumferentialdirection of flow. This circumferential flow aspect is in addition tothe general direction of the air and fuel mixture along the combustoraxis 50 from or near the surface 64 towards the transition duct 35 (seeFIG. 1). The fuel and air mixture passes through the pre-chamber 29 andinto the combustion chamber 28. The main air flow 34A forces the pilotair flow 34B and entrained fuel near to the igniter 58, which thenignites the fuel/air mixture.

To start the engine, a starter motor rotates the shaft 22, compressor 14and turbine 18 to a predetermined speed when the pilot fuel is suppliedand ignited. Once ignited the combustor internal geometry and the airflow patterns cause a pilot flame to exist. As the engine becomesself-powering the starter-motor can be switched off. As engine demand orload is increased from start-up, fuel is supplied to the main fuelinjection ports and mixed with the main air flow 34A. A main flame iscreated in the combustion chamber 28 and which is radially outwardlylocated relative to the pilot flame.

Reference is now made to FIG. 3, which shows a schematic perspective andcut-away view of part of the pilot burner 52 and in detail the liquidfuel lance 56. The liquid fuel lance 56 comprises the elongate fuellance body 86 and the liquid fuel tip 72 which are elements that can beunitary or separate. The liquid fuel tip 72 is located and captured by anarrowing 78 at an end of the first hole 66 and forms a tight fit. Atthe end of the fuel flow passage 70, the liquid fuel tip 72 includes aswirl plate 80 which defines an array of fuel conduits 82 having inletsand outlets. The fuel conduits 82, only one of which is shown, areangled relative to a longitudinal or fuel lance axis 79 of the liquidfuel lance 56. Downstream of the swirl plate 80 is a fuel swirl chamber84 and then a fuel outlet 86, which in this example is a fuel filmer.This fuel filmer 86 is divergent and produces a cone of liquid fuel. Inother examples, the fuel outlet 86 can be an orifice that produces aspray of fuel or a number of orifices, each producing a spray of fuel.

The liquid fuel tip 72 forms an array of pilot air flow conduits 88having inlets that communicate with the pilot air flow passage 69 andoutlets 90 which surround the fuel filmer 86. In this exemplaryembodiment, the pilot air flow conduits 88 are inclined or angled inboth a circumferential sense and a radially inwardly relative to thelongitudinal axis 79 of the liquid fuel lance 56. In other embodiments,the pilot air flow conduits 88 may be axially aligned, or angled in onlyone of the circumferential sense or radially inwardly relative to thelongitudinal axis 79. In this exemplary embodiment there are 8 pilot airflow conduits 88; although in other embodiments there may be more orfewer conduits.

Pilot liquid fuel flowing in the fuel flow passage 70 enters the inletsof the fuel conduits 82 and exits the outlets imparting a swirl to thefuel in the fuel swirl chamber 84. The swirling fuel forms a thin filmover the fuel filmer 86, which sheds the fuel in a relatively thin cone.Pilot air flow 34B impinges the cone of fuel and breaks the fuel intosmall droplets. The swirling vortex of air from the outlets 90 atomisesthe fuel along with the main air flow 34A.

The pilot air flow 34B is particularly useful at engine start-up and lowpower demands when the main air flow 34A has a relatively low mass flowcompared to higher power demands and because of the lower mass flow isless able to atomise the liquid fuel. Advantageously, the pilot air flow34B provides cooling to the pilot fuel lance and helps prevent fuelcoking and carbon build up on the pilot fuel lance.

FIG. 4 is a view along the combustor axis 50 and on the surface 64 ofthe burner 30 where the pilot burner 52 is generally surrounded by themain burner 54. The liquid fuel lance 56 and the igniter 58 are shownmounted in the burner body 53 of the pilot burner 52. The swirlerarrangement 55 of the main burner 54 surrounds the surface 64 anddirects the main airflow 34B via the annular array of passages 62. Theannular array of swirler vanes 60 and passages 62 are arranged to imparta tangential flow component to the main air flow 34A such that when theairflow portions from each passage 62 coalesce they form a vortex 34Cgenerally about the burner axis 50. In this embodiment, the vortex 34Crotates generally anti-clockwise as seen in FIG. 4; this vortex 34Ccould also be said to be rotating in a clockwise direction as it travelsin a direction from the surface 64 to the transition duct 35 through thepre-chamber 29 and then the combustor chamber 28.

In this exemplary embodiment, the vortex 34C is a single vortex, but inother examples the arrangements of pilot burner 52 and the main burner54 can create a number of vortices rotating in either the same directionor different directions and at different rotational speeds.

The positions of the liquid fuel lance 56 and the igniter 58 arearranged so that the swirling or rotating main air flow 34A passes overor around the liquid fuel lance 56 and then on to the igniter 58. As themain airflow forms a vortex 34C about the axis 50, the liquid fuel lance56 and the igniter 58 are positioned at approximately the same radialdistance from the axis 50. Thus as the fuel lance 56 injects or spraysliquid fuel into the pre-chamber 29 the main airflow 34C entrains thefuel and transports it towards the igniter 58, where ignition can takeplace. However, it has been found that the fuel lance 56 and the igniter58 can be located at different radial distances from the burner axis 50and preferably the igniter 58 is radially inwardly of the fuel lance 56because the co-rotating vortices draw the pilot vortex radially inwardcompared to counter-rotating vortices.

The vortex 34C has many different stream velocities within its massflow. In this example, the portion of the vortex denoted by arrow 34Csis travelling at a lower velocity than the portion of the vortex denotedby arrow 34Cf. Main air flow portion 34Cs is radially inwardly of mainair flow portion 34Cf with respect to the axis 50. Main air flow portion34Cs is at approximately the same radial position as the radially innerpart of the pilot fuel lance 56 and the main air flow portion 34Cf is atapproximately the same radial position as the radially outer part of thepilot fuel lance 56.

FIG. 5 and FIG. 6 show sectional views of the main air flow along pathsA-A and B-B respectively as shown in FIG. 4 and the distribution of fueldroplets. In FIG. 4 the flow path B-B is radially outwardly of the fuellance 56 and igniter 58 and the flow path A-A is approximately at thesame radius as at least a part of the fuel lance 56 and igniter 58.

In FIG. 6 the fuel lance 56 and igniter 58 are shown in dashed lines forreference purposes. As shown, each portion of main air flow exiting eachpassage 62 flows for a short distance immediately across the surface 64,before leaving the surface 64 and travelling away from the surface 64and along the axis 50 as another portion of the main air flow joins froma circumferentially adjacent passage 62. Thus as can be seen the anyfuel droplets 92 entrained in this portion of the main air flow longflow path B-B are quickly lifted away from the surface 64 and thereforeaway from the igniter 58.

In FIG. 5 the main air flow 34A passes over the fuel lance 56 and ontowards the igniter 58. The outlets 90, which surround the fuel filmer86 of the fuel lance 56, direct the pilot air flow 34B to impinge on thecone of fuel exiting the fuel filmer 86 and break the fuel film intosmall droplets 92. The swirling vortex of pilot air, shown schematicallyas 94, from the outlets 90 atomises the fuel as it mixes with the mainair flow 34A. The swirling vortex of pilot air 94 effectively forms afluid buffer and causes to be formed on its leeward or downstream side arecirculation zone or a low-pressure zone 96. This recirculation zone ora low-pressure zone 96 draws the main air flow 34A towards the surface64 between the fuel lance 56 and igniter 58. A portion of the fueldroplets 92 are also drawn towards the surface 64 and therefore close tothe igniter 58 such that good ignition of the fuel/air mixture ispossible.

Referring now to FIG. 7, which is a view on the tip 72 of the fuel lance56 and generally along its axis 79, the array of outlets 90 direct thepilot air flow 34B with both tangential and radial components. Thesetangential and radial components will be explained in more detail belowwith reference to FIGS. 8A and 8B. When the portions of pilot air flow34B from each outlet 90 merge they coalesce into the pilot vortex 94.The pilot vortex 94 rotates in a generally anti-clockwise direction asseen in FIG. 7; this vortex 94 could also be said to be rotating in aclockwise direction as it travels in a direction from the surface of thetip 72 towards the transition duct 35 through the pre-chamber 29 andthen the combustor chamber 28. This is the same general direction ofrotation as the main vortex. In this example, there are 8 outlets 90arranged symmetrically about the axis 79 of the fuel lance and about thefuel filmer 86. This arrangement of outlets produces, at leastinitially, a symmetric pilot vortex 94. In other examples, the outlets90 may be asymmetrically arranged about the fuel filmer 86 and one oreach of the outlets 90 may be a different size.

In the case of the known fuel lance 56 and main swirler arrangement,where oppositely rotating vortices are present, in service it has beenfound that the outlets 90 become blocked by carbon deposits formed fromliquid fuel landing on the surfaces of the fuel lance 56. In addition,carbon deposits can form on other surfaces of the burner arrangement.This blocking reduces the amount of pilot air flow 34B which in turnthis reduces the effectiveness of the pilot air flow 34B shearing andbreaking up the fuel film. As a consequence ignition of the fuel/airmixture becomes more difficult and unpredictable. Thus it has been foundthat the oppositely rotating main vortex and pilot vortex 94 causesparticular air flow characteristics that lead to liquid fuel contactingthe surface of the fuel lance and which then forms carbon deposits thatblock the outlets 90.

The counter-rotating pilot air flow 34B delivery and thecounter-rotating pilot vortex 94 remain strong enough to effectivelyform the fluid buffer 94 and cause to be formed on its leeward ordownstream side, the recirculation zone 96 or low-pressure zone 96. Thusthe recirculation zone 96 or low-pressure zone 96 still draws the mainair flow 34A towards the surface 64 between the fuel lance 56 andigniter 58. A portion of the fuel droplets 92 are also drawn towards thesurface 64 and therefore close to the igniter 58 such that good ignitionof the fuel/air mixture remains equally possible.

It has been found that arranging the fuel lance 56 and main swirlerarrangement to have their respective vortices rotating in the samerotational direction, i.e. both clockwise or both anti-clockwise, carbondeposits are prevented or substantially prevented because fewer liquiddroplets 92 contact the surfaces of the fuel lance 56 and burner.

FIG. 8A is an isometric view of the liquid fuel lance 56 schematicallyshowing the relative radial angle μ of one of the air passages 88; theother air passages have been omitted for clarity. The air assistpassages 88 have inlets 91 and a central axis 92. The air assistpassages are typically drilled, but can be laser drilled or formed by anelectron beam. It is possible that the tip of the fuel lance may beformed by layered deposition techniques, such as direct laserdeposition, and so the shape of the air assist passage can be curved inany direction and in this case the angles referred to can relate to theissues air flow direction.

The air assist passages 88 are radially angled μ relative to the fuellance axis 79. In this preferred embodiment the air assist passages 88are radially angled μ approximately 45° relative to the fuel lance axis79. However, at a minimum the angle μ is approximately 5° relative tothe fuel lance axis 79. For best results the air assist passages 88 areradially angled μ between and including 30° and 60° relative to the fuellance axis 79. It is even possible for the air assist passages 88 to beradially angled μ between and including 15° and 60° relative to the fuellance axis 79 in certain examples of the invention.

FIG. 8B is a view on the exposed surface of the tip 72 of the liquidfuel lance as seen in FIG. 3 for example. Only one of the air assistpassages 88 is shown for clarity. Here the central axis 92 of the airassist passages has a tangential angle δ of approximately 30° relativeto a tangent 93 of the fuel lance axis 79. This tangential angle δ ofapproximately +30° is relative to a tangent 93, which is to say that thecentral axis 92 is angled ‘inwardly’. Here the outlet 90 is locatedradially inwardly of the inlet 91 relative to the axis 79. By anglingthe air passage inwardly a tighter vortex is generated and which canfavourably atomise the fuel issuing from the fuel nozzle 86. In otherembodiments the tangential angle δ can be between 25 and 45° relative toa tangent 93 and this can be dependent on the spray angle or liquid coneangle of the fuel issuing from the fuel outlet.

Alternatively, the tangential angle δ of approximately −45° is relativeto a tangent 93, which is to say that the central axis 92 is angled‘outwardly’. Here the outlet 90 is located radially inwardly of theinlet 91 relative to the axis 79. This can generates a weaker or lesstight vortex to be formed and which can be more favourable where a highvolume of air is used for the air assist or where a wide or flat cone offuel is generated from the fuel orifice 86. In other examples, the airassist passages 88 can have a tangential angle δ approximately 0°relative to a tangent of the fuel lance axis 79.

Referring to FIG. 9 which is a view on the surface 64 of the burner 30and along the axis 50 and from which a radial line 102 emanates andpasses through the axis 78 of the fuel lance 56. The fuel lance 56 andigniter 58 are shown along with main airflow arrows 34A issuing from themain air flow passages 62. As described earlier, the portion of thevortex denoted by arrow 34Cf is travelling at a generally highervelocity than the portion of the vortex denoted by arrow 34Cs. Therelatively slower flow is generally radially inward of the fastervelocity air.

The fuel lance 56 as previously described is at least partly housedwithin the burner body 53 of the burner 30 and the outlets 90 and thefuel filmer 86 are located at or near to the surface 64. In thisexample, the outlets 90 and the fuel filmer 86 are located below thesurface 64 in the burner body 53. The igniter 58 is also at least partlyhoused within the burner body 53 and has an end face 59, located justbelow the surface 64, but could be at or near to the surface 64.

The burner 30 further includes an array of gas injection ports 122generally formed in a radially outward part of the burner 30 and under acircumferential lip 124 as shown in FIG. 2. These gas injection ports122 can supply a pilot gas-fuel as is known in the art.

The terms clockwise and anticlockwise are with respect to the view onthe surface 64 of the burner 30 as seen in FIG. 9.

What is claimed is:
 1. A burner for a combustor of a gas turbinecombustor, the burner comprises comprising: a body having a surface andan burner axis, a fuel lance, an igniter and a main air flow passage orpassages, wherein the main air flow passage or passages is angledrelative to the burner axis and creates a main vortex about the burneraxis in a first rotational direction, the main vortex travels in adirection along the burner axis and away from the surface, wherein thefuel lance is located downstream of the igniter with respect to thefirst rotational direction of the main vortex so that a part of a mainair flow washes over the fuel lance and then over the igniter, whereinthe fuel lance comprises a fuel lance axis, a liquid fuel tip having afuel outlet and an array of air assist passages having outlets arrangedabout the fuel outlet, wherein the air assist passages are angledrelative to the fuel lance axis to create an air assist vortex about thefuel lance axis in the same rotational direction to the first rotationaldirection.
 2. A-The burner for a combustor of a gas turbine combustor asclaimed in claim 1, wherein the main air flow passage or passages istangentially angled relative to the burner axis.
 3. The burner for acombustor of a gas turbine combustor as claimed in claim 1, wherein theair assist passages are radially angled μ relative to the fuel lanceaxis.
 4. The burner for a combustor of a gas turbine combustor asclaimed in claim 3, wherein the air assist passages are radially angledμ between and including 15° and 60° relative to the fuel lance axis. 5.The burner for a combustor of a gas turbine combustor as claimed inclaim 3, wherein the air assist passages are radially angled μapproximately 30° relative to a tangent of the fuel lance axis.
 6. Theburner for a combustor of a gas turbine combustor as claimed in claim 1,wherein the air assist passages have a tangential angle δ between +/−45°relative to a tangent of the fuel lance axis.
 7. The burner for acombustor of a gas turbine combustor as claimed in claim k wherein theair assist passages have a tangential angle δ approximately 0° relativeto a tangent of the fuel lance axis.
 8. The burner for a combustor of agas turbine combustor as claimed in claim 1, wherein the fuel lance andthe igniter are located at the same radial distance from the burneraxis.
 9. The burner for a combustor of a gas turbine combustor asclaimed in claim 1, wherein the fuel lance and the igniter are locatedat different radial distances from the burner axis.
 10. The burner for acombustor of a gas turbine combustor as claimed in claim 1, wherein thefuel outlet of the fuel lance is located at or near to the surface. 11.The burner for a combustor of a gas turbine combustor as claimed inclaim 1, wherein the igniter is at least partly housed within the bodyand has an end face, the end face is located at or near to the surface.12. The burner for a combustor of a gas turbine combustor as claimed inclaim 1, wherein the burner comprises an annular array of swirl vanesarranged about the burner axis and which form the main air flowpassages.
 13. The burner for a combustor of a gas turbine combustor asclaimed in claim 1, wherein the main air flow passages are angled in ananti-clockwise direction and the air assist passages are angled in ananti-clockwise direction relative to a view on to the surface.
 14. Theburner for a combustor of a gas turbine combustor as claimed in claim 1,wherein the main air flow passages are angled in a clockwise directionand the air assist passages are angled in a clockwise direction relativeto a view on to the surface.
 15. The burner for a combustor of a gasturbine combustor as claimed in claim 1, wherein the fuel outlet is anyone of a fuel prefilmer, an orifice or a number of orifices.
 16. Theburner for a combustor of a gas turbine combustor as claimed in claim 9,wherein the igniter is radially inwardly of the fuel lance.